CN107272653B - Fault diagnosis method for flight control system - Google Patents

Fault diagnosis method for flight control system Download PDF

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CN107272653B
CN107272653B CN201710593924.0A CN201710593924A CN107272653B CN 107272653 B CN107272653 B CN 107272653B CN 201710593924 A CN201710593924 A CN 201710593924A CN 107272653 B CN107272653 B CN 107272653B
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CN107272653A (en
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陈欣
张榕
李继广
堵湘君
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Nanjing University of Aeronautics and Astronautics
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B23/00—Testing or monitoring of control systems or parts thereof
    • G05B23/02—Electric testing or monitoring
    • G05B23/0205—Electric testing or monitoring by means of a monitoring system capable of detecting and responding to faults
    • G05B23/0208—Electric testing or monitoring by means of a monitoring system capable of detecting and responding to faults characterized by the configuration of the monitoring system
    • G05B23/0213—Modular or universal configuration of the monitoring system, e.g. monitoring system having modules that may be combined to build monitoring program; monitoring system that can be applied to legacy systems; adaptable monitoring system; using different communication protocols
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B2219/00—Program-control systems
    • G05B2219/20—Pc systems
    • G05B2219/24—Pc safety
    • G05B2219/24065—Real time diagnostics

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Abstract

The invention relates to a fault diagnosis method of a flight control system, and belongs to the technical field of fault diagnosis of flight control systems. The method comprises the following steps: step 1) determining the composition and the connection structure of a flight control system; step 2) determining the structure and redundancy configuration of the distributed flight control computer; step 3) determining an aircraft object model; step 4) designing an actuating mechanism and a model fault observer of the airborne sensing system according to the object model determined in the step 3), and identifying typical faults; step 5) configuring corresponding model fault observers in each distributed unit according to the distributed flight control computer structure; step 6) carrying out fault diagnosis on the flight control system according to the hardware redundancy and the model fault observer; and 7) testing the fault diagnosis effect of the system and optimizing the system structure. The method can realize the fault diagnosis of the flight control computer, the airborne sensing system and the executing mechanism in the flight control system.

Description

Fault diagnosis method for flight control system
Technical Field
The invention relates to a fault diagnosis method for a flight control system, and belongs to the technical field of fault diagnosis of flight control systems.
Background
The reliability and safety of flight control systems, including flight control computers, actuators, and airborne sensor systems, have been an important research direction for aviation aircraft. The flight control computer has the work responsibility of completing the flight control task of the unmanned aerial vehicle and managing and scheduling the airborne equipment, so that resource sharing and data information fusion are realized, and the flight control computer is a contact junction of an air command system and a ground command system. At present, higher requirements on the reliability and safety of a flight control computer are provided by increasing the number and complexity of tasks of an unmanned aerial vehicle, fusing multiple data of sensors, increasing the flight range and performance boundary and the like.
Each functional unit in the distributed flight control computer is independently separated, each part runs different functional programs, and complex flight control and communication tasks are separated. The redundancy technology can receive the same information by using a plurality of same functional units and modules in the design, and acquire the working state of the unit through redundancy voting and detection.
The faults of an actuating mechanism and a sensor belong to the external faults of a flight control computer, the output control of the flight control computer to the actuating mechanism and the receiving of sensor information are completed through an internal external interface of the flight control computer, the traditional single fault observer used at present has the problem of insufficient fault diagnosis coverage rate, the fault detection of the actuating mechanism and the sensor cannot be separated from the faults of the external interface of the flight control computer, and then the internal faults of the flight control computer cannot be detected and isolated.
Disclosure of Invention
Aiming at the defect of insufficient fault diagnosis coverage rate, the invention provides the fault diagnosis method for the flight control system, which can effectively diagnose the faults of a flight control computer, an actuating mechanism and a sensor in the flight control system, solves the defect that the traditional single fault observer cannot diagnose the internal faults of the computer, and improves the fault diagnosis coverage rate and reliability of the flight control system.
The invention adopts the following technical scheme for solving the technical problems:
a flight control system fault diagnosis method comprises the following steps:
step 1) determining a composition structure of a flight control system, wherein the composition structure comprises a distributed flight control computer, an execution mechanism, an airborne sensing system and a system connection structure;
step 2) determining the structure and redundancy configuration of the distributed flight control computer;
step 3) determining an aircraft object model;
step 4) designing an actuating mechanism and a model fault observer of the airborne sensing system according to the object model determined in the step 3), and identifying typical transverse deviation faults, gradual change faults and sudden change faults;
step 5) configuring corresponding model fault observers in each distributed unit according to the distributed flight control computer structure;
step 6) fault diagnosis is carried out according to the hardware redundancy and the model fault observer;
and 7) testing the fault diagnosis effect of the system and optimizing the system structure.
The distributed flight control computer in the step 2) adopts a distributed structure and comprises a three-redundancy central processing unit, a double-redundancy serial communication unit, a double-redundancy analog quantity interface unit and a double-redundancy switching value interface unit, wherein the three-redundancy central processing unit, the double-redundancy serial communication unit, the double-redundancy analog quantity interface unit and the double-redundancy switching value interface unit are respectively in bidirectional connection with an internal bus; the three-redundancy central processing unit is used for solving the control law of the whole flight control system and managing peripheral equipment; the dual-redundancy serial communication unit is responsible for communicating with external serial equipment; the dual-redundancy analog interface unit is responsible for acquiring analog quantity information and outputting analog quantity control information; the dual-redundancy switching value unit is responsible for collecting the equipment state and outputting equipment control information.
And 4) inputting the control data of the actuating mechanism and the airborne sensing data into the fault observer in the step 4), and outputting fault diagnosis results including fault types and fault estimation values.
The fault observer is divided into an actuating mechanism fault observer and an airborne sensing system fault observer, and the two observers do not have a coupling characteristic.
The specific content of configuring the corresponding model fault observer by each distributed unit in the step 5) is as follows:
the three-redundancy central processing unit and the dual-redundancy analog quantity unit are provided with an actuating mechanism fault observer; the dual-redundancy serial communication unit configures a fault observer of the on-board sensing system.
The specific method for performing fault diagnosis according to the hardware redundancy and the model fault observer in the step 6) is as follows: the input execution mechanism control data of the execution mechanism fault observer configured by the three-redundancy central processing unit are the execution mechanism control data resolved by the three-redundancy central processing unit and the airborne sensing data acquired by the dual-redundancy serial communication unit, and the output of the execution mechanism fault observer is the fault information of the three-redundancy central processing unit;
the fault diagnosis coverage surface comprises analog quantity unit A/D functional interface faults, analog quantity unit D/A functional interface faults and execution mechanism faults;
the fault observer of the airborne sensing system configured by the dual-redundancy serial communication unit inputs control data of an execution mechanism calculated by the three-redundancy central processing unit and airborne sensing data acquired by the dual-redundancy serial communication unit, outputs fault information of the dual-redundancy serial communication unit, performs fault diagnosis through four groups of information of sensor information acquired by the dual-redundancy serial communication unit and a fault observer result, and the fault diagnosis coverage surface comprises serial input/output functional interface faults of the serial communication unit and airborne sensing system faults.
The invention has the following beneficial effects:
1. and a distributed structure is adopted, so that the complexity of a software program is reduced, and the reliability of the flight control computer is improved.
2. The fault diagnosis coverage rate is improved, and the fault diagnosis of a flight control computer, an actuating mechanism and a sensor in a flight control system can be effectively carried out.
Drawings
FIG. 1 flight control system architecture.
Fig. 2 rudder circuit-actuator structure.
FIG. 3 is a distributed flight control computer architecture.
FIG. 4 is a flight control system information flow diagram.
FIG. 5 model fault observer architecture.
FIG. 6(a) is a schematic diagram of lateral deviation fault types; FIG. 6(b) is a schematic diagram of a gradual failure type; fig. 6(c) is a schematic diagram of the abrupt fault type.
Fig. 7 is a view showing a failure diagnosis structure of the CPU unit.
Fig. 8 serial communication unit external interface failure diagnosis structure.
FIG. 9 is an analog unit external interface fault diagnosis architecture.
Detailed Description
The technical solution of the present invention is further described in detail below with reference to the accompanying drawings.
The invention discloses a fault diagnosis method for a flight control system of a hybrid redundancy system comprising hardware redundancy and model observer resolution redundancy. The structure of the flight control system is shown in fig. 1, and comprises a distributed flight control computer, a sensor system and an actuating mechanism; the rudder loop-actuator structure is shown in fig. 2, and the actuator structure receives control data from a flight control computer and feeds back the response result of the actuator. The structure of the distributed flight control computer is shown in fig. 3, in the Distributed Flight Control Computer (DFCC), a central processing unit (CPU unit) adopts a triple-redundancy configuration, and a serial communication unit (SIO unit), an analog unit (AIO unit) and a switching value unit (DIO unit) adopt a double-redundancy configuration, in fig. 3, an interface between the serial communication unit and an external device is represented as TRX; the analog quantity unit D/A module is an analog quantity output control interface, and the A/D module is an analog quantity acquisition input interface; the D/O of the switching value unit is a switching value output interface, and the D/I is a switching value input interface. The three-redundancy central processing unit performs control law resolving of the whole flight control system and realizes management of peripheral equipment; the dual-redundancy serial communication unit is responsible for communicating with external serial equipment; the dual-redundancy analog interface unit is responsible for acquiring analog quantity information and outputting analog quantity control information; the dual-redundancy switching value unit is responsible for collecting the equipment state and outputting equipment control information. The analog interface unit, the switching value interface unit and the serial interface unit are data acquisition and output parts of the distributed flight control computer, and the three functional units are responsible for receiving sensor data, acquiring states of peripheral equipment and uploading the states to the central processing unit and outputting a control law resolving result and logic management data of the central processing unit.
The schematic diagram of the internal information flow of the flight control system is shown in fig. 4, the flight control computer analog quantity D/A module controls the actuating mechanism, the A/D module receives the feedback result of the actuating mechanism and analog quantity sensor data, the serial communication unit receives sensor information, some sensors simultaneously send the analog quantity data and the serial communication data, the data analysis types are the same, and the implementation of the technical scheme is not influenced.
The model fault observer structure is shown in fig. 5, with an observer within the dashed box. Typical fault types are shown in fig. 6 and are classified as lateral deviation, gradual change and abrupt change faults.
The design of the implementation of the flight control system fault diagnosis process is described in detail below.
1. Model fault observer
Let the dynamic equation of the flight control system be
Figure BDA0001355405060000041
Wherein x (t) e Rn，u(t)∈Rm，y(t)∈RpRespectively a state vector, an input vector and an output vector of the system;
Figure BDA0001355405060000051
is the differential of the state vector, representing the rate of change of the state vector; a, B and C are known dimension-adaptive matrixes, and (A and C) are observable pairs.
Faults of an actuating mechanism, a sensor and the like can be divided into additive faults and multiplicative faults, the additive faults are taken as examples in the application, fault models of the actuating mechanism and the sensor are shown in formulas (2) and (3),
Figure BDA0001355405060000052
Figure BDA0001355405060000053
in the formula (f)a(t) is the actuator fault for the system, E is the actuator fault parameter matrix, and E ═ B; f. ofsAnd (t) is the sensor fault of the system, and D is a parameter matrix fed back by the sensor fault.
The actuator control fault observer is
Figure BDA0001355405060000054
Order to
Figure BDA0001355405060000055
The observer for sensor communication fault is
Figure BDA0001355405060000056
Order to
Figure BDA0001355405060000057
In the formula (I), the compound is shown in the specification,
Figure BDA0001355405060000058
respectively a state vector and an output vector of the observer;
Figure BDA0001355405060000059
is a differential of the observed state vector, representing the rate of change of the observed state vector; the dimension-adaptive matrix is a gain matrix of the observer L;
Figure BDA00013554050600000510
observing a fault for the actuator;
Figure BDA00013554050600000511
observing a fault for the sensor; e.g. of the typex(t) is the error of the observer state from the actual state; e.g. of the typey(t) is the error between the observed output and the actual output of the system;
Figure BDA00013554050600000512
observing the error between the fault and the actual fault for the actuating mechanism;
Figure BDA00013554050600000513
the error of the fault from the actual fault is observed for the sensor.
The actuator error dynamic equation is expressed as equation (8), the sensor error dynamic equation is expressed as equation (9),
Figure BDA0001355405060000061
Figure BDA0001355405060000062
in the formula (I), the compound is shown in the specification,
Figure BDA0001355405060000063
is the rate of change of error of the observer state from the actual state.
The fault estimation algorithm of the sudden fault and the gradual fault is
Figure BDA0001355405060000064
Figure BDA0001355405060000065
In the formula (I), the compound is shown in the specification,
Figure BDA0001355405060000066
a differential, representing the rate of change of the observed fault, for the observed fault;
Figure BDA0001355405060000067
is the rate of change of the actual fault;
Figure BDA0001355405060000068
is ey(t) a differential, representing the rate of change of the error; gamma is a learning rate matrix of fault estimation; f is an adaptive matrix of fault estimates,
Figure BDA0001355405060000069
to observe the error in the rate of change of fault and the actual rate of change of fault.
CPU unit fault diagnosis
The CPU unit fault diagnosis structure is shown in FIG. 7, the CPU unit is provided with an actuator fault observer Ob _ a (U, Y), U and Y respectively represent observer control input and sensor feedback information; u in FIG. 7rControl output information resolved for the CPU unit; y isSSensor information input for a serial communication unit, YSMFor transmission to the CPU unit by the serial communication unitThe sensor information of (1).
The three-redundancy CPU unit inputs the same sensor signal, the three-redundancy control calculation and the fault observer of the actuating mechanism are carried out in the main control unit, and the fault diagnosis result is obtained according to the control data calculated by the three-redundancy CPU unit and the feedback input observer of the sensor.
3. Serial communication unit external interface fault diagnosis
The serial communication unit adopts a redundancy fault diagnosis structure formed by hardware dual-redundancy input and a sensor communication fault observer, as shown in fig. 8, the serial communication unit operates the sensor communication fault observer Ob _ s (U, Y), and a sensor signal fed back to the CPU unit is YsM. The dual-redundancy serial communication unit inputs the same sensor signal YsInput as Y via the serial interfacesA，YsBThe two form a master-slave backup mode, and the observer Ob _ s (U, Y) is used to analyze redundancy to form a four-redundancy form in the master serial communication unit according to YsA，YsBAnd Ob _ s (U, Y) results are subjected to the fault diagnosis judgment of the serial communication interface, as shown in table 1.
TABLE 1 Serial communications Unit external interface Fault diagnosis
Figure BDA0001355405060000071
As shown in table 1, when the serial interface inputs of the two groups are the same (1) to (4), and the fault model observers output fault information, the actuator is in a fault state; when the output information of the two fault model observers are different, the information of the non-fault unit is taken as YsM(ii) a When the serial interface inputs of the two groups are different (5) - (7), the unit interface fault can be judged; case (8) requires comparison with CPU unit diagnostic information to derive the faulty unit location.
4. Fault diagnosis of external interface of analog quantity unit
The structure of the analog quantity unit for diagnosing the fault is shown in FIG. 9, the CPU unit operates the executing mechanism to control the fault observer Ob _ a (U, Y), and the observer control input is used for solving the control data U for the CPU unitr. Simulation ofThe quantity unit operation actuating mechanism controls a fault observer, and the control input of the observer is feedback control data U of the actuating mechanismbAircraft sensor State is YsM. The analog quantity unit diagnoses the A/D, D/A function module fault and the actuator fault in the analog quantity unit by using an observer combination mode.
TABLE 2 analog unit external interface Fault diagnostics
Figure BDA0001355405060000081
As shown in table 2, 4 judgment results are formed according to the difference between the results of the CPU unit fault observer and the analog unit fault observer, which can effectively cover a no-fault state and three fault locations, and the fault type is estimated according to a fault estimation algorithm and fed back to the main control unit for corresponding fault adjustment.
It will be understood by those skilled in the art that, unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the prior art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein.
The above-mentioned embodiments, objects, technical solutions and advantages of the present invention are further described in detail, it should be understood that the above-mentioned embodiments are only illustrative of the present invention and are not intended to limit the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (2)

1. A flight control system fault diagnosis method is characterized by comprising the following steps:
step 1) determining a composition structure of a flight control system, wherein the composition structure comprises a distributed flight control computer, an execution mechanism, an airborne sensing system and a system connection structure;
step 2) determining the structure and redundancy configuration of the distributed flight control computer; the distributed flight control computer adopts a distributed structure and comprises a three-redundancy central processing unit, a dual-redundancy serial communication unit, a dual-redundancy analog interface unit and a dual-redundancy switching value interface unit, wherein the three-redundancy central processing unit, the dual-redundancy serial communication unit, the dual-redundancy analog interface unit and the dual-redundancy switching value interface unit are respectively in bidirectional connection with an internal bus; the three-redundancy central processing unit is used for solving the control law of the whole flight control system and managing peripheral equipment; the dual-redundancy serial communication unit is responsible for communicating with external serial equipment; the dual-redundancy analog quantity interface unit is responsible for acquiring analog quantity information and outputting analog quantity control information; the dual-redundancy switching value unit is responsible for acquiring equipment states and outputting equipment control information;
step 3) determining an aircraft object model;
step 4) designing an actuating mechanism and a model fault observer of the airborne sensing system according to the object model determined in the step 3), and identifying typical transverse deviation faults, gradual change faults and sudden change faults; the input of the fault observer is execution mechanism control data and airborne sensing data, and the output of the fault observer is a fault diagnosis result which comprises a fault type and a fault estimation value;
step 5) configuring corresponding model fault observers in each distributed unit according to the distributed flight control computer structure; the specific content of configuring the corresponding model fault observer by each distributed unit is as follows: the three-redundancy central processing unit and the dual-redundancy analog quantity unit are provided with an actuating mechanism fault observer; the dual-redundancy serial communication unit is configured with a fault observer of the airborne sensing system;
step 6) fault diagnosis is carried out according to the hardware redundancy and the model fault observer; the specific method for performing fault diagnosis according to the hardware redundancy and the model fault observer is as follows: the input execution mechanism control data of the execution mechanism fault observer configured by the three-redundancy central processing unit are the execution mechanism control data resolved by the three-redundancy central processing unit and the airborne sensing data acquired by the dual-redundancy serial communication unit, and the output of the execution mechanism fault observer is the fault information of the three-redundancy central processing unit;
the fault diagnosis coverage surface comprises analog quantity unit A/D functional interface faults, analog quantity unit D/A functional interface faults and execution mechanism faults;
the fault observer of the airborne sensing system configured by the dual-redundancy serial communication unit inputs actuating mechanism control data resolved by the three-redundancy central processing unit and airborne sensing data acquired by the dual-redundancy serial communication unit, the output of the fault observer is fault information of the dual-redundancy serial communication unit, fault diagnosis is carried out through four groups of information of sensor information acquired by the dual-redundancy serial communication unit and a fault observer result, and a fault diagnosis coverage surface comprises serial input/output functional interface faults of the serial communication unit and airborne sensing system faults;
and 7) testing the fault diagnosis effect of the system and optimizing the system structure.
2. The method for diagnosing the faults of the flight control system according to claim 1, wherein the fault observer is divided into an actuator fault observer and an airborne sensing system fault observer, and the two observers do not have coupling characteristics.
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